The assignee of the present invention manufactures and deploys spacecraft for communications and broadcast services. Many such spacecraft operate in a geosynchronous orbit having a period equal to one sidereal day (approximately 23.93 hours).
A particular type of geosynchronous orbit is a geostationary orbit (GSO), characterized as being substantially circular and co-planar with the Earth's equator. A satellite in GSO is substantially stationary relative to any point on the earth's surface. Requirements for orienting such a spacecraft and provisions for recovering from a loss of desired pointing orientation are described in Brumfield et. al., US 2007/0228218 (hereinafter, Brumfield) which claims the priority benefit of U.S. 60/776,858. The disclosures of the foregoing related applications are hereby incorporated by reference in their entirety into the present application.
An elevation angle from a user located on the Earth to a satellite in GSO is a function of the user's latitude. When a service area on the ground intended to receive communications or broadcast services (hereinafter, an “intended service area”) is at a relatively high latitude, the elevation angle is relatively small. At the latitudes of service areas containing many population centers of interest, for example in North America, Europe, and Asia, the elevation angle from the intended service area to the GSO spacecraft is small enough that service outages, for example from physical blockage, multipath fading, and foliage attenuation, are problematic.
To mitigate this problem, satellites operable in inclined, elliptical geosynchronous orbits have been proposed, as described, for example in Briskman, et al., U.S. Pat. No. 6,223,019, (hereinafter, Briskman) the disclosure of which is hereby incorporated in its entirety into the present patent application. A geosynchronous, highly inclined, elliptical orbit (HIEO) may be selected such that the orbit's apogee is located at a pre-selected, substantially constant, longitude and latitude. A satellite disposed in an HIEO can, during much of its orbital period (e.g., sixteen hours out of twenty four) enable higher elevation angles to a user than a GSO satellite.
A satellite disposed in an HIEO, however, has a substantial motion with respect to a user. This motion is illustrated in FIG. 1, which presents an exemplary “ground track” 101 of a spacecraft in a typical HIEO. Ground track 101 represents successive nadir points of the orbiting spacecraft as the spacecraft moves around an HIEO orbit track. In the example illustrated in FIG. 1, the spacecraft nadir crosses the equator at hour zero at approximately −68 degrees longitude, at which time, it is within view of a service area centered at approximately 40 degrees north latitude and −96 degrees longitude. During the next sixteen hours, the spacecraft nadir (i.e., ground track) remains within the northern hemisphere, reaching a maximum latitude at orbit apogee of about 55 degrees at a longitude of −96 degrees.
FIG. 1 illustrates that a satellite disposed in a HIEO may deliver communications or broadcast services to an intended service area for as much as sixteen hours centered about the orbit apogee (hereinafter, the “service period”). During the service period a ‘look’ angle (that may be conventionally described in terms of azimuth and elevation angle from a location in the intended service area to the spacecraft) varies considerably. At any moment during the service period, there exists an optimal spacecraft payload orientation that maximizes the equivalent isotropic radiated power (EIRP) from, for example, a transmit antenna to the intended service area. In other words, to optimize delivery of broadcast or communications services, a boresight of the antenna should be steered (relative to nadir pointing) so as to point the boresight at the coordinate location of an optimal target. Thereby, the EIRP over the intended service area may be continuously optimized during the service period.
The optimal target is, in general, not fixed with respect to the earth, the satellite, or the satellite nadir. Rather, the optimal target moves, for example, along “Z axis aim” track 102 which may be substantially different from ground track 101. As illustrated in FIG. 2, which presents region ‘A’ of FIG. 1 on an expanded scale, for an intended service area nominally centered at 39.5 degree north latitude and −96 degrees longitude, the optimal target may vary over a range between 35 and 40 degrees north latitude and between −91 and −101 degrees longitude.
The problem of payload pointing compensation for spacecraft in geosynchronous, near-geostationary orbit is discussed in Fowell, U.S. Pat. No. 6,135,389 (hereinafter, “Fowell”), the disclosure of which is hereby incorporated in its entirety into the present patent application. According to the techniques described in Fowell, deviations from a perfect geostationary orbit may be compensated by steering a satellite payload steered at a fixed virtual target below the Earth's surface. To the extent that a geosynchronous orbit is substantially inclined, this technique becomes inadequate. Even for minor inclinations, the technique described in Fowell fails to result in steering the satellite payload to track the coordinate location of the optimal target.
Payload pointing compensation for spacecraft in highly inclined elliptical orbits have been provided for in the prior art by means of offsets modeled by polynomial and Fourier series functions. In the prior art solutions, the polynomial and Fourier series function are parameterized only with respect to time and provide correction only for drift in an orbit's argument of latitude. Moreover, solutions provided by the prior art provide for steering of a spacecraft antenna reflector with respect to the spacecraft bus. See, for example, Briskman, et al., “S-DARS Broadcast from Inclined, Elliptical Orbits, IAF-01-M.5.04, 52nd International Astronautical Congress, October 2001. Such steering can have a detrimental effect on payload optics, and may be impossible for certain large, unfurlable reflectors.